CN108263639A - Aircaft configuration key position fatigue life on-line monitoring method based on indirect measuring strain under spectrum carries - Google Patents

Aircaft configuration key position fatigue life on-line monitoring method based on indirect measuring strain under spectrum carries Download PDF

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Publication number
CN108263639A
CN108263639A CN201810080596.9A CN201810080596A CN108263639A CN 108263639 A CN108263639 A CN 108263639A CN 201810080596 A CN201810080596 A CN 201810080596A CN 108263639 A CN108263639 A CN 108263639A
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fatigue
strain
monitoring
key position
life
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CN108263639B (en
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尚德广
惠杰
李道航
李志高
薛龙
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Beijing University of Technology
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Beijing University of Technology
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    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64F—GROUND OR AIRCRAFT-CARRIER-DECK INSTALLATIONS SPECIALLY ADAPTED FOR USE IN CONNECTION WITH AIRCRAFT; DESIGNING, MANUFACTURING, ASSEMBLING, CLEANING, MAINTAINING OR REPAIRING AIRCRAFT, NOT OTHERWISE PROVIDED FOR; HANDLING, TRANSPORTING, TESTING OR INSPECTING AIRCRAFT COMPONENTS, NOT OTHERWISE PROVIDED FOR
    • B64F5/00—Designing, manufacturing, assembling, cleaning, maintaining or repairing aircraft, not otherwise provided for; Handling, transporting, testing or inspecting aircraft components, not otherwise provided for
    • B64F5/60—Testing or inspecting aircraft components or systems

Abstract

Aircaft configuration key position fatigue life on-line monitoring method based on indirect measuring strain under being carried the invention discloses spectrum is related to the monitoring technical field of aircraft structure fatigue Fatigue crack initiation, steps of the method are：(1) structural member dangerous point and calibration sensor position and the factor of stress concentration are calculated using finite element method；(2) installation sensor carries out online strain monitoring；(3) strain carried out using rain flow method to monitoring carries out cycle count, and acquire the A LOCAL STRESS-STRAIN of dangerous point；(4) damage for calculating cycle adds up, and judges whether to start crack initiation monitoring；(5) crack initiation assessment parameter is calculated, comparative evaluation parameter chooses whether to continue after judging the germinating situation of fatigue crack with discriminant parameter；(6) it chooses whether to continue after investigating damage accumulated result.Monitoring result illustrates that this method can preferably monitor aircraft structure key position fatigue crack life situations.

Description

The aircaft configuration key position fatigue life based on indirect measuring strain is online under spectrum carries Monitoring method
Technical field
Application field of the present invention is fatigue life monitoring direction, refers in particular to the aircraft based on indirect measuring strain under a kind of spectrum carries Structural key position fatigue life on-line monitoring method.
Background technology
Aircraft country economic field, field of traffic, military field in occupation of critical role, as caused by aircraft fatigue Structural break can cause serious accident to occur.Aircaft configuration fracture accident is often by the germinating of a fatigue crack , when crack Propagation to certain length, can directly cause structure destruction leads to accident.Therefore, the present invention proposes one kind For the fatigue life on-line monitoring method of aircaft configuration key position, ensure that aircraft safety is reliably on active service, there is important reality Border meaning.
Current service life on-line monitoring method is typically the installation strain biography around aircraft key position stress concentration position Sensor estimates that the ess-strain of dangerous point comes mathematic(al) expectation with this, but service life one that this traditional computational methods obtain As differ 2 times or so or even error bigger with actual life, and crack monitoring generally requires more complicated sensor and too In the sensitive situation that will appear false alarm, therefore, it is superimposed on the basis of conventional method a kind of equally based on stress concentration position Surrounding strains the fatigue crack initiation monitoring method of on-line measurement, can be more nearly the practical on-line monitoring service life.It is proposed Method also for the service life of other mechanical structure key positions on-line monitoring provide one have application value technology.
Invention content
Present invention aims to meet the needs of aircraft structure fatigue service life monitoring, it is proposed that between a kind of spectrum is based under carrying The aircaft configuration key position fatigue life on-line monitoring method of measuring strain is connect, this method is also applied for monitoring other machinery knot The fatigue life of the key position of structure.
Aircaft configuration key position based on indirect measuring strain under technical solution provided by the present invention is carried for a kind of spectrum Fatigue life on-line monitoring method, step are：
Step 1)：The position of wanted monitoring of structures key position dangerous point is determined using finite element method, while demarcates and wants In the position of practical structures installation strain transducer, No. 1 sensor is mounted on the back side of dangerous point face, No. 2 sensor installations In No. 1 sensor homonymy, the actual (real) thickness of aircraft structure is needed to refer to the distance of No. 1 sensor, ensures two sensors Distance is not less than aircraft structure thickness and cannot be more than twice of aircraft structure thickness, and two sensing stations of guarantee Line is vertical with prediction crack propagation plane, and the stress concentration between this is determined using No. 2 sensors and the stress of dangerous point COEFFICIENT KtTo be used for the A LOCAL STRESS-STRAIN at subsequent calculations monitoring position；
Step 2)：According to the sensing station that step 1) determines, strain transducer is installed on aircraft structure, for reality When monitoring No. 1 sensor and No. 2 sensors strain value, be denoted as ε respectively1And ε2；
Step 3)：A load ε in the block of survey is taken into custody using rain flow method docking2Time history carries out cycle meter Number, while using the cylic stress-strain curve of the structural material, the A LOCAL STRESS-STRAIN of dangerous point is determined according to Neuber methods；
Step 4)：It is calculated using Smith formula by rain-flow counting to the load block ε of reception2It is extracted in time history The corresponding fatigue damage D of i-th of cyclei, it is shown below,
σmax--- the maximum stress of the cycle；
The range of strain of Δ ε --- the cycle；
σ′f--- fatigue strength coefficient；
ε′f--- tired plastic coefficient；
E --- Young's modulus；
Ni--- i-th of cycle corresponding service life；
B --- fatigue strength exponent；
C --- tired plasticity index；
Total damage D of received load block is accumulated using Miner theorems, wherein
When always damage D accumulations are to 0.5, continue in next step, as D ＜ 0.5, to return to step 3) and continue to acquire next load Block carries out cycle calculations；N represents cycle total degree.
Step 5)：Calculating receives ε in load block1And ε2The absolute value is obtained with load block in the absolute value of maximum value difference The derivative of the increased curve of number assesses parameter K as fatigue crack initiation, when fatigue crack initiation assessment parameter K is more than to differentiate to join Number KcWhen, show that aircaft configuration key position has fatigue crack initiation, i.e. fatigue crack is formed, immediately warning stop using or into Row Measuring error, if being less than discriminant parameter KcThen continue step 6).
Step 6)：Total damage D accumulation situations are investigated, as D >=0.9, characterize the remaining life of prediction less than 10%, it is alert Close to the limit, if selection is continuing with return to step 3) continues to calculate report prompting life consumption, otherwise stops using aircraft And Measuring error, as D ＜ 0.9, return to step 3) and continue to calculate.
The loading conditions i.e. loading spectrum of aircraft structure to be monitored is set to a load block during one flight of aircraft is risen and fallen.
The discriminant parameter is carried out in advance by finite element modelling actual crack extension or using simulating piece fatigue test Calibration.
Compared with prior art, the present invention has the advantages that.
The advantage of the invention is that：It is tired to propose the aircaft configuration key position based on indirect measuring strain under a kind of spectrum carries Labor service life on-line monitoring method.The strain transducer that this method is utilized need not be mounted on aircraft structure dangerous point position, and It is to investigate the strain variation around dangerous point by installing strain transducer in dangerous point position other side, is come with this indirect Reflect the fatigue crack initiation at dangerous point position and fatigue life situation, therefore the size of pair of strain sensors and type be not tight The limitation of lattice can also select suitable strain transducer as needed under harsh working environment, be more conducive to be applied to each The real-time monitoring of fatigue crack is carried out under kind environment, for example, the environment such as high temperature, fuel oil.And this method is in traditional life prediction The monitoring method of the fatigue crack initiation life based on same set of strain transducer is increased in the method for calculating, in this way can either The forewarning function of classic fatigue life-span prediction method is played, and the false alarm that can evade the fatigue crack initiation monitoring of early period shows As ensure that the accuracy and reliability of service life on-line monitoring.
Description of the drawings
Fig. 1 the method for the present invention realizes the flow chart of fatigue crack on-line monitoring.
The strain transducer scheme of installation of Fig. 2 the method for the present invention.
Fig. 3 the method for the present invention is applied to the fatigue crack monitoring effect figure of certain aircraft structure.
Specific embodiment
The specific embodiment of the present invention is described with reference to the drawings.
The present invention is further illustrated the present invention by aircraft structure fatigue test,
Aircaft configuration key position fatigue life on-line monitoring method based on indirect measuring strain under a kind of spectrum carries, specifically Computational methods are as follows：
Step 1)：The position of this structural member key position dangerous point is determined using finite element method, while calibration will be in reality The position of border structure installation strain transducer, No. 1 sensor are mounted on the back side of dangerous point face, and No. 2 sensors are mounted on No. 1 Sensor homonymy needs to refer to the actual (real) thickness of aircraft structure with the distance of No. 1 sensor, ensures two sensor distances not Less than aircraft structure thickness and cannot be more than twice of aircraft structure thickness, and ensure the lines of two sensing stations with It predicts that crack propagation plane is vertical, the stress concentration factor K between this is determined using No. 2 sensors and the stress of dangerous pointt To be used for the A LOCAL STRESS-STRAIN at subsequent calculations position；
Step 2)：According to the sensing station that step 1) determines, strain transducer is installed on the aircraft structure, is used for The strain value of No. 1 and No. 2 sensor of monitoring in real time, is denoted as ε respectively1And ε2；
Step 3)：Using rain flow method to the load block ε of reception2Cycle count is carried out, while uses the structural material Cylic stress-strain curve, the A LOCAL STRESS-STRAIN of dangerous point is determined according to Neuber methods；
Step 4)：I-th of cycle corresponding service life is calculated using Smith formula, it is as follows,
σmax--- the maximum stress of the cycle；
The plastic strain amplitude of Δ ε --- the cycle；
σ′f--- fatigue strength coefficient；
ε′f--- tired plastic coefficient；
E --- Young's modulus；
Ni--- i-th of cycle corresponding service life；
B --- fatigue strength exponent；
C --- tired plasticity index；
It is accumulated using Miner theorems to always damaging D, wherein
When always damage D accumulations are to 0.5, continue in next step, as D ＜ 0.5, to return to step 3) and continue cycling through calculating；
Step 5)：Calculating receives ε in load block1And ε2The value is obtained with the increased song of load block number in the absolute value of difference The derivative of line assesses parameter K as fatigue crack initiation, when fatigue crack initiation assessment parameter K is more than discriminant parameter Kc(this is sentenced Other parameter can be demarcated by finite element modelling actual crack extension or by simulating piece fatigue test in advance) when, characterize aircraft There is fatigue crack initiation at structural key position, and fatigue to longevity, warning stops using aircraft to be detected repair immediately, if do not surpassed Discriminant parameter Kc is crossed then to continue in next step.
Step 6)：Total damage D accumulation situations are investigated, as D >=0.9, the remaining life of theory of representation prediction is insufficient 10%, close to the limit, if selection is continuing with return to step 3) continues to calculate alarm sounds life consumption, otherwise stops making With aircraft and Measuring error, as D ＜ 0.9, return to step 3) and continue to calculate..
The advantage of the invention is that：It is tired to propose the aircaft configuration key position based on indirect measuring strain under a kind of spectrum carries Labor service life on-line monitoring method.The strain transducer that this method is utilized need not be mounted on aircraft structure dangerous point position, and It is to investigate the strain variation around dangerous point by installing strain transducer in dangerous point position other side, is come with this indirect Reflect the fatigue crack initiation at dangerous point position and fatigue life situation, therefore the size of pair of strain sensors and type be not tight The limitation of lattice can also select suitable strain transducer as needed under harsh working environment, be more conducive to be applied to each The real-time monitoring of fatigue crack is carried out under kind environment, for example, the environment such as high temperature, fuel oil.And this method calculates in traditional service life Method on increase the monitoring method of the fatigue crack initiation based on same set of strain transducer, classic fatigue can either be played The forewarning function of life-span prediction method, and the false alarm phenomenon of the fatigue crack initiation monitoring of early period can be evaded, it ensure that the longevity Order the accuracy of on-line monitoring.
In order to verify the aircaft configuration key position fatigue life proposed by the present invention measured indirectly based on strain transducer The effect of on-line monitoring method, by the result of the monitoring obtained by this method compared with the crack initiation situation that actual observation measures, As shown in Figure 2.The result shows that after total damage D of monitoring is more than 0.9, do not find that crackle is continuing with, when the fatigue of monitoring When crack initiation assessment parameter K is more than discriminant parameter 0.1, stops using, observed fatigue crack initiation at this time, length is 2.22mm corresponds to the service life as 2875 load blocks (flight is risen and fallen), real-time theoretical calculation fatigue crack initiation life about 2176 Load block (service life when crackle just germinates, service life when corresponding total damage D accumulations are to 1), illustrates that this method captures in time The germinating of fatigue crack has accurately achieved the purpose that the real time on-line monitoring service life, while it is pre- also to remain the classic fatigue service life The forewarning function of survey method, it is therefore proposed that spectrum carry under the aircaft configuration key position fatigue life based on indirect measuring strain On-line monitoring method can accurately monitor fatigue crack initiation life situation.

Claims (3)

  1. The aircaft configuration key position fatigue life on-line monitoring method based on indirect measuring strain under 1. spectrum carries, feature exist In：The implementation steps of this method are as follows,
    Step 1)：The position of wanted monitoring of structures key position dangerous point is determined using finite element method, while calibration will be in reality The position of border structure installation strain transducer, No. 1 sensor are mounted on the back side of dangerous point face, and No. 2 sensors are mounted on No. 1 Sensor homonymy needs to refer to the actual (real) thickness of aircraft structure with the distance of No. 1 sensor, ensures two sensor distances not Less than aircraft structure thickness and cannot be more than twice of aircraft structure thickness, and ensure the lines of two sensing stations with It predicts that crack propagation plane is vertical, the stress concentration factor K between this is determined using No. 2 sensors and the stress of dangerous pointt To be used for the A LOCAL STRESS-STRAIN at subsequent calculations monitoring position；
    Step 2)：According to the sensing station that step 1) determines, strain transducer is installed on aircraft structure, for supervising in real time The strain value of No. 1 sensor and No. 2 sensors is surveyed, is denoted as ε respectively1And ε2；
    Step 3)：A load ε in the block of survey is taken into custody using rain flow method docking2Time history carries out cycle count, simultaneously Using the cylic stress-strain curve of the structural material, the A LOCAL STRESS-STRAIN of dangerous point is determined according to Neuber methods；
    Step 4)：It is calculated using Smith formula by rain-flow counting to the load block ε of reception2I-th extracted in time history Recycle corresponding fatigue damage Di, it is shown below,
    σmax--- the maximum stress of the cycle；
    The range of strain of Δ ε --- the cycle；
    σ′f--- fatigue strength coefficient；
    ε′f--- tired plastic coefficient；
    E --- Young's modulus；
    Ni--- i-th of cycle corresponding service life；
    B --- fatigue strength exponent；
    C --- tired plasticity index；
    Total damage D of received load block is accumulated using Miner theorems, wherein
    When always damage D accumulations are to 0.5, continue in next step, as D ＜ 0.5, return to step 3) continue to acquire next load block into Row cycle calculations；N represents cycle total degree；
    Step 5)：Calculating receives ε in load block1And ε2The absolute value of maximum value difference is obtained the absolute value and increases with load block number The derivative of the curve added assesses parameter K as fatigue crack initiation, when fatigue crack initiation assessment parameter K is more than discriminant parameter Kc When, show that aircaft configuration key position has fatigue crack initiation, i.e. fatigue crack is formed, and warning immediately is stopped using or examined Repair is surveyed, if being less than discriminant parameter KcThen continue step 6)；
    Step 6)：Total damage D accumulation situations are investigated, as D >=0.9, characterize the remaining life of prediction less than 10%, alarm carries Show life consumption close to the limit, returning to step 3) if selection is continuing with continues to calculate, and otherwise stops using aircraft and examines Repair is surveyed, as D ＜ 0.9, step 3) is returned to and continues to calculate.
  2. The aircaft configuration key position fatigue life based on indirect measuring strain is online under 2. spectrum according to claim 1 carries Monitoring method, it is characterised in that：The loading conditions i.e. loading spectrum of aircraft structure to be monitored is determined during one flight of aircraft is risen and fallen For a load block.
  3. The aircaft configuration key position fatigue life based on indirect measuring strain is online under 3. spectrum according to claim 1 carries Monitoring method, it is characterised in that：The discriminant parameter is in advance by finite element modelling actual crack extension or using simulating piece Fatigue test is demarcated.
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CN109733641A (en) * 2019-01-19 2019-05-10 北京工业大学 A kind of aircraft full size structure part multiaxle fatigue experimental method
CN110704979A (en) * 2019-09-30 2020-01-17 三一海洋重工有限公司 Crane life evaluation method and device and electronic equipment
CN110861784A (en) * 2019-12-04 2020-03-06 中国直升机设计研究所 Calculation and monitoring method for fatigue accumulation damage of helicopter flight test load
CN111191391A (en) * 2019-12-10 2020-05-22 中国航空工业集团公司成都飞机设计研究所 Simulation piece design method based on local stress distribution weighting coefficient
CN111289160A (en) * 2020-02-13 2020-06-16 中国特种设备检测研究院 Stress state monitoring system and method for large-scale amusement facility
CN112487580A (en) * 2020-11-27 2021-03-12 苏州热工研究院有限公司 Method for evaluating running state of important pump gear box of nuclear power plant
CN112520064A (en) * 2020-12-04 2021-03-19 中国航空工业集团公司成都飞机设计研究所 Automatic damage identification method based on strain monitoring
CN112685937A (en) * 2020-12-25 2021-04-20 中国航空工业集团公司沈阳飞机设计研究所 Airplane structure health monitoring method based on fiber bragg grating sensor
CN114506472A (en) * 2022-02-23 2022-05-17 贵州贵飞飞机设计研究院有限公司 Method for evaluating stress at key point of structural stress
CN116046980A (en) * 2022-12-16 2023-05-02 北京航空航天大学 Structural fatigue damage diagnosis method based on strain monitoring
CN117262237A (en) * 2023-11-22 2023-12-22 北京航空航天大学 Aircraft cockpit lid skeleton simulation piece fatigue test method considering assembly dispersibility

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Cited By (18)

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Publication number Priority date Publication date Assignee Title
CN109733641A (en) * 2019-01-19 2019-05-10 北京工业大学 A kind of aircraft full size structure part multiaxle fatigue experimental method
CN110704979A (en) * 2019-09-30 2020-01-17 三一海洋重工有限公司 Crane life evaluation method and device and electronic equipment
CN110861784A (en) * 2019-12-04 2020-03-06 中国直升机设计研究所 Calculation and monitoring method for fatigue accumulation damage of helicopter flight test load
CN111191391A (en) * 2019-12-10 2020-05-22 中国航空工业集团公司成都飞机设计研究所 Simulation piece design method based on local stress distribution weighting coefficient
CN111191391B (en) * 2019-12-10 2022-08-23 中国航空工业集团公司成都飞机设计研究所 Simulation piece design method based on local stress distribution weighting coefficient
CN111289160B (en) * 2020-02-13 2022-01-04 中国特种设备检测研究院 Stress state monitoring system and method for large-scale amusement facility
CN111289160A (en) * 2020-02-13 2020-06-16 中国特种设备检测研究院 Stress state monitoring system and method for large-scale amusement facility
CN112487580A (en) * 2020-11-27 2021-03-12 苏州热工研究院有限公司 Method for evaluating running state of important pump gear box of nuclear power plant
CN112487580B (en) * 2020-11-27 2023-11-24 苏州热工研究院有限公司 Nuclear power plant important pump gear box operational state evaluation method
CN112520064A (en) * 2020-12-04 2021-03-19 中国航空工业集团公司成都飞机设计研究所 Automatic damage identification method based on strain monitoring
CN112685937A (en) * 2020-12-25 2021-04-20 中国航空工业集团公司沈阳飞机设计研究所 Airplane structure health monitoring method based on fiber bragg grating sensor
CN112685937B (en) * 2020-12-25 2022-09-20 中国航空工业集团公司沈阳飞机设计研究所 Airplane structure health monitoring method based on fiber grating sensor
CN114506472A (en) * 2022-02-23 2022-05-17 贵州贵飞飞机设计研究院有限公司 Method for evaluating stress at key point of structural stress
CN114506472B (en) * 2022-02-23 2023-09-12 中航贵州飞机有限责任公司 Evaluation method for stress at structural stress key point
CN116046980A (en) * 2022-12-16 2023-05-02 北京航空航天大学 Structural fatigue damage diagnosis method based on strain monitoring
CN116046980B (en) * 2022-12-16 2023-09-01 北京航空航天大学 Structural fatigue damage diagnosis method based on strain monitoring
CN117262237A (en) * 2023-11-22 2023-12-22 北京航空航天大学 Aircraft cockpit lid skeleton simulation piece fatigue test method considering assembly dispersibility
CN117262237B (en) * 2023-11-22 2024-02-02 北京航空航天大学 Aircraft cockpit lid skeleton simulation piece fatigue test method considering assembly dispersibility

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